System and method for locating a probe within a gas turbine engine

ABSTRACT

A method for locating probes within a gas turbine engine may generally include positioning a plurality of location transmitters relative to the engine and inserting a probe through an access port of the engine, wherein the probe includes a probe tip and a location signal receiver configured to receive location-related signals transmitted from the location transmitters. The method may also include determining a current location of the probe tip within the engine based at least in part on the location-related signals and identifying a virtual location of the probe tip within a three-dimensional model of the engine corresponding to the current location of the probe tip within the engine. Moreover, the method may include providing for display the three-dimensional model of the engine, wherein the virtual location of the probe tip is displayed as a visual indicator within the three-dimensional model.

FIELD OF THE INVENTION

The present subject matter relates generally to gas turbine engines and,more particularly, to a system and method for locating a probe within agas turbine engine.

BACKGROUND OF THE INVENTION

A gas turbine engine typically includes a turbomachinery core having ahigh pressure compressor, combustor, and high pressure turbine in serialflow relationship. The core is operable in a known manner to generate aprimary gas flow. The high pressure compressor includes annular arrays(“rows”) of stationary vanes that direct air entering the engine intodownstream, rotating blades of the compressor. Collectively one row ofcompressor vanes and one row of compressor blades make up a “stage” ofthe compressor. Similarly, the high pressure turbine includes annularrows of stationary nozzle vanes that direct the gases exiting thecombustor into downstream, rotating blades of the turbine. Collectivelyone row of nozzle vanes and one row of turbine blades make up a “stage”of the turbine. Typically, both the compressor and turbine include aplurality of successive stages.

In order to allow for periodic inspection of the core parts of theengine (e.g., the compressor blades and the turbine blades), borescopeports are typically provided in the engine casings and/or frames. Suchports allow optical borescope instruments to be inserted into the coreengine to enable a visual inspection of the engine to be performedwithout requiring disassembly of the engine components. However, once aninstrument has been inserted into a borescope port, minimal informationis typically available to an inspector regarding the actual position ofthe instrument within the engine, leading to errors in measurements andreducing the efficiency of performing the visual inspection.

Accordingly, a system and method for locating a probe relative to a gasturbine engine as such probe is being inserted within the engine wouldbe welcomed within the technology.

BRIEF DESCRIPTION OF THE INVENTION

Aspects and advantages of the invention will be set forth in part in thefollowing description, or may be obvious from the description, or may belearned through practice of the invention.

In one aspect, the present subject matter is directed to a method forlocating probes within a gas turbine engine. The method may generallyinclude positioning a plurality of location transmitters relative to thegas turbine engine, wherein each location transmitter has a knownlocation relative to the gas turbine engine. The method may also includeinserting a probe through an access port of the gas turbine engine,wherein the probe includes a probe tip and a location signal receiverconfigured to receive location-related signals transmitted from thelocation transmitters. In addition, the method may include receiving thelocation-related signals at a computing device communicatively coupledto the probe, determining, by the computing device, a current locationof the probe tip within the gas turbine engine based on thelocation-related signals and the known locations of the locationtransmitters and identifying, with the computing device, a virtuallocation of the probe tip within a three-dimensional model of the gasturbine engine corresponding to the current location of the probe tipwithin the gas turbine engine. Moreover, the method may includeproviding for display, by the computing device, the three-dimensionalmodel of the gas turbine engine, wherein the virtual location of theprobe tip is displayed as a visual indicator within thethree-dimensional model.

In another aspect, the present subject matter is directed to a systemfor locating probes within a gas turbine engine. The system maygenerally include a plurality of location transmitters, wherein eachlocation transmitter is positioned at a known location relative to thegas turbine engine. The system may also include a probe configured to beinserted through an access port of the gas turbine engine. The probe mayinclude a probe tip and a location signal receiver configured to receivelocation-related signals transmitted from the location transmitters. Inaddition, the system may include a computing device communicativelycoupled to the probe. The computing device may be configured todetermine a current location of the probe tip within the gas turbineengine based on the location-related signals received by the locationsignal receiver and the known locations of the plurality of locationtransmitters. Moreover, the computing device may be configured toidentify a virtual location of the probe tip within a three-dimensionalmodel of the gas turbine engine corresponding to the current location ofthe probe tip within the gas turbine engine and provide for display thethree-dimensional model of the gas turbine engine, wherein the virtuallocation of the probe tip is displayed as a visual indicator within thethree-dimensional model.

These and other features, aspects and advantages of the presentinvention will be better understood with reference to the followingdescription and appended claims. The accompanying drawings, which areincorporated in and constitute a part of this specification, illustrateembodiments of the invention and, together with the description, serveto explain the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present invention, including thebest mode thereof, directed to one of ordinary skill in the art, is setforth in the specification, which makes reference to the appendedfigures, in which:

FIG. 1 illustrates a cross-sectional view of one embodiment of a gasturbine engine that may be utilized within an aircraft in accordancewith aspects of the present subject matter;

FIG. 2 illustrates a partial, cross-sectional view of one embodiment ofa turbine suitable for use within the gas turbine engine shown in FIG.1, particularly illustrating access ports defined in the engine forproviding internal access to the turbine;

FIG. 3 illustrates a partial, cross-sectional view of one embodiment ofa compressor suitable for use within the gas turbine engine shown inFIG. 1, particularly illustrating access ports defined in the engine forproviding internal access to the compressor;

FIG. 4 illustrates a simplified view of one embodiment of an opticalprobe that may be used in accordance with aspects of the present subjectmatter to visually inspect a gas turbine engine;

FIG. 5 illustrates a simplified, schematic view of one embodiment of asystem for locating a probe within a gas turbine engine in accordancewith aspects of the present subject matter;

FIG. 6 illustrates a simplified view of a user interface that may bepresented to a user of the disclosed system in accordance with aspectsof the present subject matter, particularly illustrating a visualindicator being used to represent a virtual location of a probe within athree-dimensional model of a gas turbine engine that corresponds to thecurrent real-world location of the probe within the engine; and

FIG. 7 illustrates a flow diagram of one embodiment of a method forlocating a probe within a gas turbine engine in accordance with aspectsof the present subject matter.

DETAILED DESCRIPTION OF THE INVENTION

Reference now will be made in detail to embodiments of the invention,one or more examples of which are illustrated in the drawings. Eachexample is provided by way of explanation of the invention, notlimitation of the invention. In fact, it will be apparent to thoseskilled in the art that various modifications and variations can be madein the present invention without departing from the scope or spirit ofthe invention. For instance, features illustrated or described as partof one embodiment can be used with another embodiment to yield a stillfurther embodiment. Thus, it is intended that the present inventioncovers such modifications and variations as come within the scope of theappended claims and their equivalents.

In general, the present subject matter is directed to a system andmethod for locating a probe within a gas turbine engine. Specifically,in several embodiments, an optical probe used to visually inspect a gasturbine engine may be equipped with a suitable receiver configured toreceive location-related signals from a plurality of locationtransmitters installed relative to the engine (e.g., within access portsof the engine). The location-related signals received by the probereceiver may then be transmitted to a computing device configured todetermine the current location of the optical probe within the gasturbine engine, such as by using a trilateration or a triangulationtechnique to calculate the location of the probe based on thelocation-related signals and the known locations of the transmitters. Inaddition, the computing device may be configured to correlate thedetermined real-world location of the optical probe to a virtuallocation within a three-dimensional model of the gas turbine engine. Thethree-dimensional model may then be presented to a user of the systemwith the current location of the optical probe being displayed withinthe model (e.g., as a symbol or other visual indicator), which may allowthe user to more accurately position the probe within the gas turbineengine as a visual inspection of the engine is being performed. Inaddition, by determining the location of the optical probe within theengine, the computing device may also be configured to identify theinternal component(s) located adjacent to the probe tip. The name(s)and/or reference number(s) of the adjacent internal component(s) maythen be presented to the user of the system along with the currentlocation of the optical probe.

It should be appreciated that, although the present subject matter willgenerally be described herein with reference to determining the locationof an optical probe within a gas turbine engine, the disclosed systemand method may be generally used to determine the location of any probeinserted within a gas turbine engine. For instance, the system andmethod may be used to determine the location of a repair probe that hasbeen inserted within a gas turbine engine via one of its access ports toallow for a given repair procedure to be performed within the engine.

It should also be appreciated that the disclosed system and method maygenerally be used to locate probes inserted within any suitable type ofgas turbine engine, including aircraft-based turbine engines andland-based turbine engines, regardless of the engine's current assemblystate (e.g., fully or partially assembled). Additionally, with referenceto aircraft engines, it should be appreciated that the present subjectmatter may be used on wing or off wing.

Referring now to the drawings, FIG. 1 illustrates a cross-sectional viewof one embodiment of a gas turbine engine 10 that may be utilized withinan aircraft in accordance with aspects of the present subject matter,with the engine 10 being shown having a longitudinal or axial centerlineaxis 12 extending therethrough for reference purposes. In general, theengine 10 may include a core gas turbine engine (indicated generally byreference character 14) and a fan section 16 positioned upstreamthereof. The core engine 14 may generally include a substantiallytubular outer casing 18 that defines an annular inlet 20. In addition,the outer casing 18 may further enclose and support a booster compressor22 for increasing the pressure of the air that enters the core engine 14to a first pressure level. A high pressure, multi-stage, axial-flowcompressor 24 may then receive the pressurized air from the boostercompressor 22 and further increase the pressure of such air. Thepressurized air exiting the high-pressure compressor 24 may then flow toa combustor 26 within which fuel is injected into the flow ofpressurized air, with the resulting mixture being combusted within thecombustor 26. The high energy combustion products are directed from thecombustor 26 along the hot gas path of the engine 10 to a first (highpressure) turbine 28 for driving the high pressure compressor 24 via afirst (high pressure) drive shaft 30, and then to a second (lowpressure) turbine 32 for driving the booster compressor 22 and fansection 16 via a second (low pressure) drive shaft 34 that is generallycoaxial with first drive shaft 30. After driving each of turbines 28 and32, the combustion products may be expelled from the core engine 14 viaan exhaust nozzle 36 to provide propulsive jet thrust.

Additionally, as shown in FIG. 1, the fan section 16 of the engine 10may generally include a rotatable, axial-flow fan rotor assembly 38 thatis configured to be surrounded by an annular fan casing 40. It should beappreciated by those of ordinary skill in the art that the fan casing 40may be configured to be supported relative to the core engine 14 by aplurality of substantially radially-extending, circumferentially-spacedoutlet guide vanes 42. As such, the fan casing 40 may enclose the fanrotor assembly 38 and its corresponding fan rotor blades 44. Moreover, adownstream section 46 of the fan casing 40 may extend over an outerportion of the core engine 14 so as to define a secondary, or by-pass,airflow conduit 48 that provides additional propulsive jet thrust.

It should be appreciated that, in several embodiments, the second (lowpressure) drive shaft 34 may be directly coupled to the fan rotorassembly 38 to provide a direct-drive configuration. Alternatively, thesecond drive shaft 34 may be coupled to the fan rotor assembly 38 via aspeed reduction device 37 (e.g., a reduction gear or gearbox) to providean indirect-drive or geared drive configuration. Such a speed reductiondevice(s) may also be provided between any other suitable shafts and/orspools within the engine 10 as desired or required.

During operation of the engine 10, it should be appreciated that aninitial air flow (indicated by arrow 50) may enter the engine 10 throughan associated inlet 52 of the fan casing 40. The air flow 50 then passesthrough the fan blades 44 and splits into a first compressed air flow(indicated by arrow 54) that moves through conduit 48 and a secondcompressed air flow (indicated by arrow 56) which enters the boostercompressor 22. The pressure of the second compressed air flow 56 is thenincreased and enters the high pressure compressor 24 (as indicated byarrow 58). After mixing with fuel and being combusted within thecombustor 26, the combustion products 60 exit the combustor 26 and flowthrough the first turbine 28. Thereafter, the combustion products 60flow through the second turbine 32 and exit the exhaust nozzle 36 toprovide thrust for the engine 10.

The gas turbine engine 10 may also include a plurality of access portsdefined through its casings and/or frames for providing access to theinterior of the core engine 14. For instance, as shown in FIG. 1, theengine 10 may include a plurality of access ports 62 (only six of whichare shown) defined through the outer casing 18 for providing internalaccess to one or both of the compressors 22, 24 and/or for providinginternal access to one or both of the turbines 28, 32. In severalembodiments, the access ports 62 may be spaced apart axially along thecore engine 14. For instance, the access ports 62 may be spaced apartaxially along each compressor 22, 24 and/or each turbine 28, 32 suchthat at least one access port 62 is located at each compressor stageand/or each turbine stage for providing access to the internalcomponents located at such stage(s). In addition, the access ports 62may also be spaced apart circumferentially around the core engine 14.For instance, a plurality of access ports 62 may be spaced apartcircumferentially around each compressor stage and/or turbine stage.

It should be appreciated that, although the access ports 62 aregenerally described herein with reference to providing internal accessto one or both of the compressors 22, 24 and/or for providing internalaccess to one or both of the turbines 28, 32, the gas turbine engine 10may include access ports 62 providing access to any suitable internallocation of the engine 10, such as by including access ports 62 thatprovide access within the combustor 26 and/or any other suitablecomponent of the engine 10.

Referring now to FIG. 2, a partial, cross-sectional view of the first(or high pressure) turbine 28 described above with reference to FIG. 1is illustrated in accordance with embodiments of the present subjectmatter. As shown, the first turbine 28 may include a first stage turbinenozzle 66 and an annular array of rotating turbine blades 68 (one ofwhich is shown) located immediately downstream of the nozzle 66. Thenozzle 66 may generally be defined by an annular flow channel thatincludes a plurality of radially-extending, circularly-spaced nozzlevanes 70 (one of which is shown). The vanes 70 may be supported betweena number of arcuate outer bands 72 and arcuate inner bands 74.Additionally, the circumferentially spaced turbine blades 68 maygenerally be configured to extend radially outwardly from a rotor disk(not shown) that rotates about the centerline axis 12 (FIG. 1) of theengine 10. Moreover, a turbine shroud 76 may be positioned immediatelyadjacent to the radially outer tips of the turbine blades 68 so as todefine the outer radial flowpath boundary for the combustion products 60flowing through the turbine 28 along the hot gas path of the engine 10.

As indicated above, the turbine 28 may generally include any number ofturbine stages, with each stage including an annular array of nozzlevanes and follow-up turbine blades 68. For example, as shown in FIG. 2,an annular array of nozzle vanes 78 of a second stage of the turbine 28may be located immediately downstream of the turbine blades 68 of thefirst stage of the turbine 28.

Moreover, as shown in FIG. 2, a plurality of access ports 62 may bedefined through the turbine casing and/or frame, with each access port62 being configured to provide access to the interior of the turbine 28at a different axial location. Specifically, as indicated above, theaccess ports 62 may, in several embodiments, be spaced apart axiallysuch that each access port 62 is aligned with or otherwise providesinterior access to a different stage of the turbine 28. For instance, asshown in FIG. 2, a first access port 62A may be defined through theturbine casing/frame to provide access to the first stage of the turbine28 while a second access port 62B may be defined through the turbinecasing/frame to provide access to the second stage of the turbine 28.

It should be appreciated that similar access ports 62 may also beprovided for any other stages of the turbine 28 and/or for any turbinestages of the second (or low pressure) turbine 32. It should also beappreciated that, in addition to the axially spaced access ports 62shown in FIG. 2, access ports 62 may be also provided at differingcircumferentially spaced locations. For instance, in one embodiment, aplurality of circumferentially spaced access ports may be definedthrough the turbine casing/frame at each turbine stage to provideinterior access to the turbine 28 at multiple circumferential locationsaround the turbine stage.

Referring now to FIG. 3, a partial, cross-sectional view of the highpressure compressor 24 described above with reference to FIG. 1 isillustrated in accordance with embodiments of the present subjectmatter. As shown, the compressor 24 may include a plurality ofcompressor stages, with each stage including both an annular array offixed compressor vanes 80 (only one of which is shown for each stage)and an annular array of rotatable compressor blades 82 (only one ofwhich is shown for each stage). Each row of compressor vanes 80 isgenerally configured to direct air flowing through the compressor 24 tothe row of compressor blades 82 immediately downstream thereof.

Moreover, the compressor 24 may include a plurality of access ports 62defined through the compressor casing/frame, with each access port 62being configured to provide access to the interior of the compressor 24at a different axial location. Specifically, in several embodiments, theaccess ports 62 may be spaced apart axially such that each access port62 is aligned with or otherwise provides interior access to a differentstage of the compressor 24. For instance, as shown in FIG. 3, first,second, third and fourth access ports 62 a, 62 b, 62 c, 62 d areillustrated that provide access to four successive stages, respectively,of the compressor 24.

It should be appreciated that similar access ports 62 may also beprovided for any of the other stages of the compressor 24 and/or for anyof the stages of the low pressure compressor 22. It should also beappreciated that, in addition to the axially spaced access ports 62shown in FIG. 3, access ports 62 may be also provided at differingcircumferentially spaced locations. For instance, in one embodiment, aplurality of circumferentially spaced access ports may be definedthrough the compressor casing/frame at each compressor stage to provideinterior access to the compressor 24 at multiple circumferentiallocations around the compressor stage.

Referring now to FIG. 4, a simplified view of one embodiment of a probe100 that may be utilized to perform a visual inspection of a gas turbineengine 10 is illustrated in accordance with aspects of the presentsubject matter. As shown, the probe 100 has been inserted through anaccess port 62 of the engine 10, such as any of the access ports 62described above with reference to FIGS. 1-3.

In general, the probe 100 may correspond to any suitable probeconfigured to be inserted within the gas turbine engine 10 via an accessport 62. Specifically, as shown in the illustrated embodiment, the probe100 corresponds to an optical probe 100. In such an embodiment, theoptical probe 100 may correspond to any suitable optical device that maybe inserted through an access port 62 of the gas turbine engine 10 toallow images of the interior of the engine 10 to be captured orotherwise obtained. For instance, in several embodiments, the opticalprobe 100 may correspond to a borescope, videoscope, fiberscope or anyother similar optical device known in the art that allows for theinterior of a gas turbine engine 10 to be viewed through an access port62. In such embodiments, the optical probe 100 may include one or moreoptical elements (indicated schematically by dashed box 102), such asone or more optical lenses, optical fibers, image capture devices (e.g.,video cameras, still-image cameras, CCD devices, CMOS devices), cables,and/or the like, for obtaining views or images of the interior of theengine 10 at a tip 104 of the probe 100 and for transmitting or relayingsuch images from the probe tip 104 along the length of the probe 100 tothe exterior of the engine 10. For instance, as shown in FIG. 4, theinterior views or images obtained by the probe 100 may be transmittedfrom the probe tip 104 to a computing device 202 connected or otherwisecoupled to the probe 100. Additionally, as shown in FIG. 4, in oneembodiment, a light source (indicated by dashed box 106), such as anLED, may be provided at or adjacent to the probe tip 104 to providelighting within the interior of the engine 10.

The optical probe 100 may also include an articulation assembly 108 thatallows the orientation of the probe tip 104 to be adjusted within theinterior of the gas turbine engine 10. For example, the articulationassembly 108 may allow for the probe tip 104 to be rotated or pivotedabout a single axis or multiples axes to adjust the orientation of thetip 104 relative to the remainder of the probe 100. It should beappreciated that the articulation assembly 108 may generally have anysuitable configuration and/or may include any suitable components thatallow for adjustment of the orientation of the probe tip 104 relative tothe remainder of the probe 100. For example, in one embodiment, aplurality of articulation cables 110 may be coupled between the probetip 104 and one or more articulation motors 112. In such an embodiment,by adjusting the tension of the cables 110 via the motor(s) 112, theprobe tip 104 may be reoriented within the gas turbine engine 10.

It should also be appreciated that, in several embodiments, thearticulation assembly 108 may be configured to be electronicallycontrolled. Specifically, as shown in FIG. 4, the computing device 202may be communicatively coupled to the articulation assembly 108 to allowthe computing device 202 to adjust the orientation of the probe tip 104via control of the articulation assembly 108. For instance, in theillustrated embodiment, the computing device 202 may be configured totransmit suitable control signals to the articulation motor(s) 112 inorder to adjust the tension within the associated cable(s) 110, therebyallowing the computing device 202 to automatically adjust theorientation of the probe tip 104 within the gas turbine engine 10.

Additionally, as shown in FIG. 4, the optical probe 100 may also includea location signal receiver (indicated schematically by dashed box 120)positioned at or adjacent to the probe tip 104. As will be described ingreater detail below, the location signal receiver 120 may be configuredto receive location-related signals from a plurality of locationtransmitters 204 (FIG. 5) that provide an indication of the position ofthe location signal receiver 120 (and, thus, the probe tip 104) relativeto the location transmitters 204. For instance, the location signalreceiver 120 may be configured to receive signals from the locationtransmitters 204 that provide an indication of the distance definedbetween the receiver 120 and each transmitter 204 (e.g., based on thesignal strength, the time of flight of the signals and/or time ofarrival of the signals) and/or that provide an indication of the angledefined between the receiver 120 and each transmitter 204 (e.g., basedon the angle of incidence or angle of arrival of the signals). Thesignals received by the location signal receiver 120 may then betransmitted to the computing device 202 to allow the computing device202 to determine the current location of the probe tip 104 within thegas turbine engine 10 using any suitable signal-based positioningtechnique, such as a trilateration technique or a triangulationtechnique.

It should be appreciated that, in other embodiments, the probe 100 maycorrespond to any other suitable probe configured to be inserted withinthe gas turbine engine 10 via one of its access ports 62. For instance,in an alternative embodiment, the probe 100 may correspond to a repairprobe configured to be inserted within the gas turbine engine 10 toallow a repair procedure to be performed on one or more of the internalengine components, such as a probe used to repair cracks and/or otherdamage within the engine.

Referring now to FIG. 5, a simplified, schematic view of one embodimentof a system 200 for locating a probe within a gas turbine engine isillustrated in accordance with aspects of the present subject matter. Asshown, the system 200 may generally include a computing device 202, aprobe 100 installed within an access port 62 of the gas turbine engine10 and a plurality of location transmitters 204 positioned relative tothe gas turbine engine 10. As indicated above, the probe 100 may, inseveral embodiments, be configured to provide an internal view or imageof the gas turbine engine 10, which may then be transmitted to thecomputing device 202 as image data for subsequent storage thereon and/orfor presentation to a user of the system 200 via a display device 206associated with the computing device 202. Additionally, as the probe 100is being used to obtain internal images of the gas turbine engine 10,the probe 100 may also be configured to receive location-related signals(e.g., via its location signal receiver 120) transmitted from thelocation transmitters 204 that provide an indication of the position ofeach transmitter 204 relative to the probe 100. As described above, suchsignals may then be transmitted to the computing device 202 to allow thecomputing device 202 to determine the current three-dimensional locationof the probe tip 104 within the gas turbine engine 10. The determinedreal-world location of the probe tip 104 may then be displayed to theuser of the system 200 (e.g., via the display device 206) at acorresponding virtual location within a three-dimensional computer-basedmodel of the gas turbine engine 10 to provide the user with a visualindication of the current location of the probe 100 within the engine10.

In general, the computing device 202 may correspond to any suitableprocessor-based device and/or any suitable combination ofprocessor-based devices. Thus, in several embodiments, the computingdevice 202 may include one or more processor(s) 208 and associatedmemory device(s) 210 configured to perform a variety ofcomputer-implemented functions (e.g., performing the methods, steps,calculations and the like disclosed herein). As used herein, the term“processor” refers not only to integrated circuits referred to in theart as being included in a computer, but also refers to a controller, amicrocontroller, a microcomputer, a programmable logic controller (PLC),an application specific integrated circuit, and other programmablecircuits. Additionally, the memory device(s) 210 may generally comprisememory element(s) including, but not limited to, computer readablemedium (e.g., random access memory (RAM)), computer readablenon-volatile medium (e.g., a flash memory), a floppy disk, a compactdisc-read only memory (CD-ROM), a magneto-optical disk (MOD), a digitalversatile disc (DVD) and/or other suitable memory elements. Such memorydevice(s) 210 may generally be configured to store suitablecomputer-readable instructions that, when implemented by theprocessor(s) 208, configure the computing device 202 to perform variousfunctions including, but not limited to, determining the currentlocation of the probe tip 104 within the gas turbine engine 10 based atleast in part on the signals received from the transmitters 204 andproviding for display a visual indication of the location of the probetip 104 within a three-dimensional model of the engine 10.

As shown in FIG. 5, the computing device 202 may be communicativelycoupled to the probe 100 (e.g., via a communicative link or cable 212).As such, image data associated with the internal views or imagesobtained by the probe 100 may be transmitted to the computing device202. Such image data may then be used to allow the interior of the gasturbine engine 10 to be visually inspected at or adjacent to thelocation of the access port 62 within which the probe 100 has beeninserted. For example, in one embodiment, the image data may be storedwithin the device's memory 210 to allow the images to be analyzed at alater time/date to identify defects and/or damage within the engine 10.In addition (or an alternative thereto), the image data may betransmitted from the computing device 202 to the associated displaydevice 206 to allow a user of the system 200 to view the variousinternal images provided by the probe 100.

In addition, the connection provided between the probe 100 and thecomputing device 202 may also allow the location-related signalsreceived by the location signal receiver 120 from the transmitters 204to be transmitted to the computing device 202 for subsequent processingand/or analysis. For example, the transmitters 204 may be configured tocontinuously transmit location-related signals to the receiver 120 asthe probe 100 is being used to visually inspect the interior of the gasturbine engine 10. Such location-related signals may then be transmittedto the computing device 202 to allow for a real-time determination to bemade of the current location of the probe 100 within the engine 10.

In several embodiments, the location transmitters 204 may be configuredto be installed within and/or inserted through separate access ports 62of the gas turbine engine 10. For example, as shown in FIG. 5, thelocation transmitters 204 are positioned within differing access ports62 than the access port 62 within which the probe 100 is installed. Assuch, the position of each location transmitter 204 relative to the gasturbine 10 may be determined or known by identifying the specific accessport 62 within which each transmitter 204 is installed. Such knownpositions of the location transmitters 204 may then be input into and/orstored with the memory 210 of the computing device 202 for subsequentuse in determining the current location of the probe tip 104 within thegas turbine engine 10. Alternatively, the location transmitters 204 maybe configured to be positioned at any other suitable location relativeto the gas turbine engine 10 (e.g., any suitable location on, withinand/or outside of the engine 10) that allows the location-relatedsignals generated by each transmitter 204 to be received by the locationsignal receiver 120 of the probe 100. In such an embodiment, based onthe installation locations of the transmitters 204, the position of eachlocation transmitter 204 relative to gas turbine engine 10 may bedetermined and subsequently input into and/or stored within the memory208 of the computing device 202.

As shown in FIG. 5, the system 200 includes three location transmitters204 positioned relative to the gas turbine engine 10 for transmittinglocation-related signals to the location signal receiver 120 of theprobe 100. However, in other embodiments, any other suitable number oflocation transmitters 204 may be utilized within the system 200 assumingthat a sufficient amount of transmitters 204 are providing for allowingthe current location of the probe tip 104 to be determined based on thelocation-related signals transmitted by the transmitters 204 and theknown locations of such transmitters 204.

It should be appreciated that the computing device 202 may generally beconfigured to determine the current location of the probe tip 104 usingany suitable signal-based positioning technique known in the art. Forinstance, in one embodiment, the computing device 202 may be configuredto utilize a trilateration technique to determine the current locationof the probe tip 104. In such an embodiment, the computing device 202may be configured to determine the distance defined between eachlocation transmitter 204 and the probe tip 104 based on thelocation-related signals received at the location signal receiver 120.For instance, the distance between each location transmitter 204 and theprobe tip 104 may be determined based on the signal strength of thelocation-related signals received at the receiver 120 from eachtransmitter 204 (e.g., using a received signal strength indicator(RSSI)). Alternatively, the distance may be determined based on the timeof flight or time of arrival of the location-related signals receivedfrom each transmitter 204. Based on the distance defined between theprobe tip 104 and each location transmitter 204, the computing device202 may be configured to determine the three-dimensional location of theprobe tip 104 relative to the transmitters 204. Thereafter, based on theknown locations of the transmitters 204 relative to the gas turbineengine 10, the computing device 202 may calculate the currentthree-dimensional location of the probe tip 104 within the engine 10(e.g., based on the known dimensions of the gas turbine engine 10).

In another embodiment, the computing device 202 may be configured toutilize a triangulation technique to determine the current location ofthe probe tip 104. In such an embodiment, the computing device 202 maybe configured to determine the relative angle defined between eachlocation transmitter 204 and the probe tip 104 based on thelocation-related signals received at the location signal receiver 120.For instance, the receiver 120 may include a dual-antenna array thatallows the receiver 120 to detect the angle of incidence or angle ofarrival of the location-related signals received from each transmitter204. Based on the relative angle defined between the probe tip 104 andeach location transmitter 204, the computing device 202 may beconfigured to determine the three-dimensional location of the probe tip104 relative to the transmitters 204. Thereafter, based on the knownlocations of the transmitters 204 relative to the gas turbine engine 10,the computing device 202 may calculate the current three-dimensionallocation of the probe tip 104 within the engine 10 (e.g., based on theknown dimensions of the gas turbine engine 10).

Additionally, as indicated above, the computing device 202 may also beconfigured to provide for display (e.g., via the display device 206) athree-dimensional model of the gas turbine engine 10 that provides avirtual representation of the current location of the probe tip 104within the engine 10. For instance, a computer-aided design (CAD) modelor other suitable computer-based three-dimensional model may be storedwithin the memory 210 of the computing device 202. In such anembodiment, the computing device 202 may be configured to access thethree-dimensional model and compare the determined real-world locationof the probe tip 104 within the gas turbine engine 10 to thecomputer-based model so as to identify a corresponding virtual locationof the probe tip 104 within the model. The three-dimensional model maythen be presented to a user of the system 200 with the virtual locationof the probe tip 104 being displayed as a symbol or other visualindicator (e.g., a dot, star, “X” or any other suitable identifyingmark) within the model.

For example, FIG. 6 illustrates a simplified view of a user interface230 that may be presented to a user of the system 200. As shown, thethree-dimensional model of the gas turbine 10 (indicated schematicallyas cylinder 232 in FIG. 6) may be presented on the display device 206.In addition, a visual indicator 234 may be displayed within thethree-dimensional model 232 that represents the current location of theprobe tip 104 within the engine 10. As such, by manipulating thethree-dimensional model 232 (e.g., by zooming the model 232 in or out,by rotating the model 232, by making one or more of the modeledcomponents of the engine transparent and/or the like), a user of thesystem 200 may be allowed to view the exact location of the probe tip104 relative to one or more internal components of the engine 10, suchas one or more compressor blades 82 or turbine blades 68 of the engine10. The user may then manipulate the real-world location of the probetip 104 within the engine 10, as is necessary or desired, based on thevirtual location displayed within the model 232 to precisely positionthe probe tip 104 relative to the internal component(s) of the engine10.

It should be appreciated that the virtual location of the probe tip 104displayed within the three-dimensional model may be continuously updatedas the real-world location of the probe tip 104 is adjusted within thegas turbine engine 10. For example, as the user manipulates the actuallocation of the probe tip 104 within the engine 10, an updated locationfor the probe tip 104 may be determined by the computing device 202. Thevirtual location within the three-dimensional model may then be adjustedbased on the updated location of the probe tip 104, thereby providingthe user with a real-time virtual representation of the current locationof the probe tip 104.

It should also be appreciated that, in addition to the virtual locationof the probe tip 104, the computing device 202 may also be configured todisplay the name(s), part number(s) and/or other identifying informationrelated to the internal engine component(s) disposed adjacent to theprobe tip 104. For instance, as shown in FIG. 6, identifying informationassociated with an adjacent engine component(s) 236 may be presentedwithin the user interface that is being displayed to the user of thesystem 200 via the display device 206.

Additionally, it should be appreciated that, although the computingdevice 202 and the display device 206 are shown as separate componentswithin the system 200, such components may be integrated into orotherwise form part of the probe 100. For instance, in one embodiment,the computing device 202 and associated display 206 may be built intothe probe 100 as an integrated assembly.

Referring now to FIG. 7, a flow diagram of one embodiment of a method300 for locating a probe within a gas turbine engine is illustrated inaccordance with aspects of the present subject matter. In general, themethod 300 will be discussed herein with reference to the gas turbineengine 10 and the system 200 described above with reference to FIGS.1-6. However, it should be appreciated by those of ordinary skill in theart that the disclosed method 300 may generally be implemented with gasturbine engines having any other suitable engine configuration and/orwith systems having any other suitable system configuration. Inaddition, although FIG. 7 depicts steps performed in a particular orderfor purposes of illustration and discussion, the methods discussedherein are not limited to any particular order or arrangement. Oneskilled in the art, using the disclosures provided herein, willappreciate that various steps of the methods disclosed herein can beomitted, rearranged, combined, and/or adapted in various ways withoutdeviating from the scope of the present disclosure.

As shown in FIG. 7, at (302), the method 300 may include positioning aplurality of location transmitters relative to the gas turbine. Forexample, as indicated above, each location transmitter 204 may, in oneembodiment, be installed or positioned within an access port 62 of theengine 10. Alternatively, the location transmitters 204 may bepositioned at any other suitable location relative to the engine 10 thatallows for the transmitters 204 to function as described herein.

Additionally, at (304), the method 300 may include inserting a probethrough an access port of the engine. For example, as indicated above,the probe 100 may be inserted through one of the access ports 62 of theengine 10 to allow internal views or images to be obtained for visuallyinspecting the interior of the engine 10. In addition, once insertedwithin the interior of the engine 10, the probe 100 may be configured toreceive the location-related signals transmitted by the various locationtransmitters 204 (e.g., via the probe's location signal receiver 120).The location-related signals may then, at (306), be transmitted to andreceived at a computing device coupled to the probe. For example, asindicated above, the probe 100 may be communicatively coupled to acomputing device 202 to allow both the internal images obtained by theprobe 100 and the location-related signals received by the probe 100 tobe transmitted to the computing device 202.

Moreover, at (308), the method 300 may include determining a currentlocation of the probe tip within the engine based on thelocation-related signals and the known locations of the locationtransmitters. For example, in one embodiment, the computing device 202may be configured to utilize the location-related signals to calculate adistance defined between the probe tip 104 and each the locationtransmitter 204. The calculated distances along with the known locationsof the location transmitters 204 may then be used to implement atrilateration technique to determine the current location of the probetip 104 with the engine 10. In other embodiments, the computing device202 may be configured to utilize the location-related signals tocalculate an angle defined between the probe tip 104 and each thelocation transmitter 204. The calculated angles along with the knownlocations of the location transmitters 204 may then be used to implementa triangulation technique to determine the current location of the probetip 104 with the engine 10.

Referring still to FIG. 7, at (310), the method 300 may includeidentifying a virtual location of the probe tip within athree-dimensional model of the engine corresponding to the currentlocation of the probe tip within the engine. For example, as indicatedabove, the computing device 202 may be configured to correlate thereal-world location of the probe tip 104 within the engine 10 to acorresponding virtual location within the model of the engine 10.Thereafter, at (312), the computing device 202 may be configured toprovide the three-dimensional model for display with the virtuallocation of the probe tip 104 being displayed or represented as a visualindicator within the model. For instance, as indicated above, thecomputing device 202 may be configured to transmit the model to anassociated display device 206 for presentation to a user of the system200.

This written description uses examples to disclose the invention,including the best mode, and also to enable any person skilled in theart to practice the invention, including making and using any devices orsystems and performing any incorporated methods. The patentable scope ofthe invention is defined by the claims, and may include other examplesthat occur to those skilled in the art. Such other examples are intendedto be within the scope of the claims if they include structural elementsthat do not differ from the literal language of the claims, or if theyinclude equivalent structural elements with insubstantial differencesfrom the literal languages of the claims.

What is claimed is:
 1. A system for locating probes within a gas turbineengine, the system comprising: a plurality of location transmitters,each location transmitter of the plurality of location transmittersbeing respectively positioned within an access port of the gas turbineengine such that each location transmitter of the plurality of locationtransmitters are at a known location relative to the gas turbine engine;a probe inserted through an access port of the gas turbine engine thatis different than the access ports into which the location transmittersof the plurality of location transmitters are positioned, the probeincluding a probe tip and a location signal receiver, the locationsignal receiver being configured to receive location-related signalstransmitted from the plurality of location transmitters; and a computingdevice communicatively coupled to the probe, the computing device beingconfigured to: determine a current location of the probe tip within thegas turbine engine based on the location-related signals received by thelocation signal receiver and the known locations of the plurality oflocation transmitters; identify a virtual location of the probe tipwithin a three-dimensional model of the gas turbine engine correspondingto the current location of the probe tip within the gas turbine engine;and provide for display the three-dimensional model of the gas turbineengine, wherein the virtual location of the probe tip is displayed as avisual indicator within the three-dimensional model.
 2. The system ofclaim 1, wherein the plurality of location transmitters comprises atleast three location transmitters positioned relative to the gas turbineengine.
 3. The system of claim 1, wherein the computing device isconfigured to determine a distance defined between the probe tip andeach location transmitter of the plurality of location transmittersbased on the location-related signals.
 4. The system of claim 3, whereinthe computing device is further configured to determine the currentlocation of the probe tip within the gas turbine engine using atrilateration technique based on the distances defined between the probetip and the plurality of location transmitters and the known locationsof the plurality of location transmitters.
 5. The system of claim 1,wherein the computing device is configured to determine a respectiveangle defined between the probe tip and each location transmitter of theplurality of location transmitters based on the location-relatedsignals.
 6. The system of claim 5, wherein the computing device isfurther configured to determine the current location of the probe tipwithin the gas turbine engine using a triangulation technique based oneach respective angle and the known locations of the plurality oflocation transmitters.
 7. The system of claim 1, wherein the computingdevice is further configured to determine an updated location of theprobe tip within the gas turbine engine as the probe is being movedwithin the gas turbine engine.
 8. The system of claim 7, wherein thecomputing device is configured to adjust a position of the visualindicator within the three-dimensional model based on the updatedlocation of the probe tip.
 9. The system of claim 1, wherein the probecorresponds to one of a borescope, a videoscope or a fiberscope.
 10. Thesystem of claim 1, wherein the probe includes a light source forilluminating an interior of the gas turbine engine.
 11. The system ofclaim 1, wherein the probe is configured to perform a repair procedureon an internal engine component of the gas turbine engine.
 12. Thesystem of claim 1, wherein the computing device is further configured toreceive image data associated with a plurality of images obtained by theprobe of an interior of the gas turbine engine.
 13. The system of claim12, wherein the image data is configured to be inspected to identify anydefects or damage within the gas turbine engine.
 14. The system of claim12, wherein the computing device is configured to provide for displayboth the three-dimensional model of the gas turbine engine and the imagedata obtained from the probe.
 15. The system of claim 1, wherein theaccess port through which the probe is inserted is defined through aportion of a compressor of the gas turbine engine.
 16. The system ofclaim 1, wherein at least one of the plurality of access ports throughwhich the plurality of location transmitters are respectively positionedwithin is defined through a portion of a turbine of the gas turbineengine.
 17. The system of claim 1, wherein the computing device isconfigured to identify an internal component of the gas turbine enginelocated adjacent to the probe tip.